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US7074012B2 - Turbine blade - Google Patents

Turbine blade Download PDF

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Publication number
US7074012B2
US7074012B2 US10/743,091 US74309103A US7074012B2 US 7074012 B2 US7074012 B2 US 7074012B2 US 74309103 A US74309103 A US 74309103A US 7074012 B2 US7074012 B2 US 7074012B2
Authority
US
United States
Prior art keywords
face
engagement
engagement member
platform
integrally molded
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Expired - Lifetime, expires
Application number
US10/743,091
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English (en)
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US20040247442A1 (en
Inventor
Keiji Nishimura
Takahiro Ogi
Hideyuki Nishi
Toshiyuki Matsumoto
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
IHI Corp
Original Assignee
IHI Corp
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by IHI Corp filed Critical IHI Corp
Assigned to ISHIKAWAJIMA-HARIMA HEAVY INDUSTRIES CO., LTD. reassignment ISHIKAWAJIMA-HARIMA HEAVY INDUSTRIES CO., LTD. ASSIGNMENT OF ASSIGNORS INTEREST (SEE DOCUMENT FOR DETAILS). Assignors: MATSUMOTO, TOSHIYUKI, NISHI, HIDEYUKI, NISHIMURA, KEIJI, OGI, TAKAHIRO
Publication of US20040247442A1 publication Critical patent/US20040247442A1/en
Application granted granted Critical
Publication of US7074012B2 publication Critical patent/US7074012B2/en
Adjusted expiration legal-status Critical
Expired - Lifetime legal-status Critical Current

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Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/30—Fixing blades to rotors; Blade roots ; Blade spacers
    • F01D5/3007—Fixing blades to rotors; Blade roots ; Blade spacers of axial insertion type
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/10—Manufacture by removing material
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2230/00—Manufacture
    • F05D2230/60—Assembly methods
    • F05D2230/68—Assembly methods using auxiliary equipment for lifting or holding

Definitions

  • the present invention relates to a turbine blade to be installed into a female dovetail of a turbine disk of an aircraft engine.
  • a typical turbine blade includes a blade airfoil as a blade base, one side of the blade airfoil being a convex suction surface and the other side of the blade being a concave pressure surface.
  • a platform is integrally molded on the hub side (at the base end portion) of the blade and recesses are formed respectively on both sides of the platform.
  • a front seal fin protruding forward is formed at the front end of the platform and a rear seal fin protruding backward is formed at the back end of the platform.
  • a male dovetail is integrally disposed on the hub side (at the base end portion) of the platform, the dovetail has an engagement portion able to engage with a female dovetail of a turbine disk, and the engagement portion is usually formed by grinding, whereby, a jig is used for grinding, and one side of the platform can be engaged against a platform-locating portion of the jig.
  • the manufacturing process of the typical turbine blade will be described below.
  • the greater part of the turbine blade with the engagement portion remaining unfinished (an unfinished turbine blade) is molded by casting.
  • the unfinished turbine blade is located in the jig so as to make the dovetail axial direction perpendicular to the repulsive force due to work-resistance during grinding, by letting the pressure surface of the blade airfoil be supported with a support portion of the jig.
  • setting of the unfinished turbine blade onto the jig is completed by pressing the pressure surface of the blade airfoil against the location portion by means of a clamp of the jig.
  • the turbine blade is finished by forming the engagement portion along the dovetail axial direction by grinding.
  • both sides of the platform are angled to the dovetail axial direction of the engagement portion. Consequently, during formation of the engagement portion along the dovetail axial direction by grinding, a component of a force that may cause a displacement of the blade airfoil from the jig is generated. Therefore, there is a problem in that a machining tolerance of the engagement portion is degraded, lowering the quality of the turbine blade, because the unfinished turbine blade is displaced from the jig owing to an increase in the magnitude of the component of a force during grinding.
  • the unfinished turbine blade is never displaced from the jig, forming an engagement portion with tight machining tolerance, thus enhancing the quality of the turbine blade.
  • a turbine blade to be installed into an engaged member of a turbine disk of an aircraft engine is characterized in that it comprises a blade airfoil, one side of which having a convex suction surface and the other side having a concave pressure surface; a platform integrally molded on the hub side of the blade wherein a recess is formed on one side of the platform; a front seal fin formed protruding forward at the front end of the platform; and a rear seal fin formed protruding backward at the back end of the platform, an engagement member integrally molded on the hub side of the platform wherein the engagement member has a engagement face which is able to be engaged with the engaged member of turbine disk and is formed by grinding, a front engagement member integrally molded in the vicinity of a base portion of the front seal fin wherein the front engagement member has a front engagement face able to engage with a front locating portion of a jig to be used for the grinding, and the front engagement face located back from
  • the turbine blade is characterized in that the front engagement face and the rear engagement face are respectively configured to be substantially parallel to the longitudinal direction of the engagement member.
  • the turbine blade is further characterized in that the spacing between the front edge of the front engagement face and the rear edge of the rear engagement face are configured to be longer than the longitudinal length of the engagement member.
  • the turbine blade is further characterized in that the depth that the front engagement face is located back from the virtual plane and the depth that the rear engagement face is located back from the virtual plane are respectively configured to be in a range of less than or equal to 0.7 mm.
  • FIG. 1 shows a turbine blade according to an embodiment of the present invention
  • FIG. 2 is an enlarged view of the arrowed portion II in FIG. 1 ;
  • FIG. 3 shows a state where the turbine blade according to the embodiment of the present invention is set on a jig
  • FIG. 4 is a schematic view of the arrowed portion IV in FIG. 3 ;
  • FIG. 5 shows a state where the turbine blade according to the embodiment of the present invention is installed into a female dovetail of a turbine disk.
  • FIG. 1 shows a turbine blade according to an embodiment of the present invention
  • FIG. 2 is an enlarged view of the arrowed portion II in FIG. 1
  • FIG. 3 shows a state where the turbine blade according to the embodiment of the present invention is set on a jig
  • FIG. 4 is a schematic view of the arrowed portion IV in FIG. 3
  • FIG. 5 shows a state where the turbine blade according to the embodiment of the present invention is installed into a female dovetail of a turbine disk.
  • “front and rear (or back)” refers to the right hand side and left hand side in FIG. 1 and FIG. 2 , and refers to the left hand side and right hand side in FIG. 4 .
  • the turbine blade 1 relating to the embodiment of the present invention is one to be installed into a female dovetail 5 of a turbine disk 3 of a low-pressure turbine for an aircraft engine and comprises a blade airfoil 7 as a main body of the turbine blade 1 .
  • One side (the front side in FIG. 1 ) of the blade airfoil 7 is a convex suction surface 7 fa and the other side (the back side in FIG. 1 ) of the blade airfoil 7 is a concave pressure surface 7 fb.
  • a shroud 9 is integrally molded on the tip side (the outer end portion, the upside in FIG. 1 ) of the blade airfoil 7 and the shroud 9 has a couple of seal fins 11 , 13 .
  • a platform 15 is integrally molded on the hub side (the inner end portion, the downside in FIG. 1 ) of the blade airfoil 7 , and recesses 17 , 19 each having a face are formed respectively on both sides (the one side and the other side) the platform 15 . Moreover, a front seal fin 21 protruding forward is formed at the front end of the platform, and a rear seal fin 23 protruding backward is formed at the back end of the platform 15 . A so-called shank portion is also included in the platform.
  • a male dovetail 25 as an engagement member is integrally molded on the hub side of the platform 15 and the male dovetail 25 has an engagement groove (an engagement face) 25 s which is able to be engaged with an engaged protrusion (an engaged portion) 5 b of the female dovetail 5 as an engaged member, and the engagement groove 25 s is formed by grinding.
  • a front engagement member 27 is integrally molded within the recess 17 in the vicinity of a base portion of the front seal fin 21 and the front engagement member 27 has a planar front engagement face 27 f . Further, as shown with diagonal lines in FIG. 2 , a thin front semi-circular shaped wall Wf surrounding a front side-edge portion of the front engagement member 27 is integrally molded in the vicinity of the base portion of the front seal fin 21 .
  • a rear engagement member 29 is integrally molded within the recess 17 in the vicinity of a base portion of the rear seal fin 23 and the rear engagement member 29 has a planar rear engagement face 29 f .
  • a thin rear wall semi-circular shaped Wr which surrounds a rear side-edge portion of the rear engagement member 29 , is integrally molded in the vicinity of the base portion of the rear seal fin 23 .
  • the front engagement face 27 f of the front engagement member 27 is able to be engaged by a front locator pin 33 of a jig 31 to be used for the grinding
  • the rear engagement face 29 f of the rear engagement member 29 is able to be engaged against a rear locator pin 35 of the jig 31
  • the front engagement face 27 f of the front engagement member 27 and the rear engagement face 29 f of the rear engagement member 29 are respectively configured to be located slightly back from a virtual plane VF including one side of the platform 15 and also to be substantially parallel to the dovetail axial direction (longitudinal direction) of the male dovetail 25 .
  • the front engagement face 27 f is offset forward from the face of the recess toward the virtual plane VF.
  • the front engagement face 27 f is located in a first plane positioned back from the virtual plane VF.
  • the rear engagement face 29 f is located in a second plane, different from the first plane, back from the virtual plane VF and offset forward from the face of the recess toward the virtual plane VF.
  • the distance (depth) of which the front engagement face 27 f is located back from the virtual plane VF and the distance (depth) of which the rear engagement face 29 f is located back from the virtual plane VF are respectively configured to be in a range of less than or equal to 0.7 mm.
  • each of the front engagement face 27 f of the front engagement member 27 and the rear engagement face 29 f of the rear engagement member 29 has a recess in a range of less than or equal to 0.7 mm. Further, the spacing between the front edge of the front engagement face 27 f and the rear edge of the rear engagement face 29 f are configured to be longer than the length of the male dovetail 25 in the dovetail axial direction.
  • An end face of the front wall Wf and an end face of the rear wall Wr are respectively configured to be coplanar with the virtual plane VF.
  • the jig 31 includes the front locator pin 33 and a rear locator pin 35 as well as a locating roller 37 for locating the suction surface 7 fa near the tip of the blade airfoil 7 , a locating pin 41 for clipping the male dovetail 25 at the rear side thereof and a clip 39 for clipping the male dovetail 25 at the front side thereof, and a clamp 45 for pressing the pressure surface 7 fb near the hub of the blade airfoil 7 downward via a rubber pad 43 , an engagement roller 47 able to be engaged against the back end of the shroud 9 , a contact bolt 49 able to be contacted with the front end of the shroud 9 .
  • the operation (mainly manufacturing of the turbine blade 1 ) of an embodiment of the present invention will be described below.
  • the greater part of the turbine blade 1 with machining portions of the engagement groove 25 s and the shroud 9 remaining unfinished (an unfinished turbine blade 1 ′), is molded by casting. Since the vicinity of base portion of the front seal fin 21 is configured to be thicker in consideration of the strength of the fin, the front engagement member 27 and the front wall Wf are molded utilizing the thicker portion. Also, since the vicinity of base portion of the rear seal fin 23 is configured to be thicker in consideration of the strength of the fin, the rear engagement member 29 and the rear wall Wr are molded utilizing the thicker portion.
  • each of the front engagement face 27 f and the rear engagement face 29 f has a recess in a range of less than or equal to 0.7 mm, casting defects do not easily occur in the vicinity of the front engagement face 27 f and the rear engagement face 29 f .
  • the machining portion of the shroud 9 is operated by appropriate machining.
  • the front engagement face 27 f of the front engagement member 27 and the rear engagement face 29 f of the rear engagement member 29 are engaged by the front locator pin 33 of the jig 31 and the rear locator pin 35 of the jig 31 respectively, and the suction surface 7 fa of the blade airfoil 7 is made to be located with the locating roller 37 of the jig 31 .
  • the unfinished turbine blade 1 ′ can be located in the jig 31 so as to make the dovetail axial direction perpendicular to the repulsive force due to work-resistance during grinding.
  • location of the unfinished turbine blade 1 ′ relative to the jig 31 in the machining direction is also performed by engaging the back end of the shroud 9 with the engagement roller 47 of the jig 31 to make the contact bolt 49 contact with the front end of the shroud 9 .
  • the male dovetail 25 is located at the rear end of the dovetail 25 with a locating pin 41 of the jig 31 , and clipped with a clip 39 at the front end of the dovetail 25 , and the pressure surface 7 fb near the hub of the blade airfoil 7 is pressed downward with a clamp 45 of the jig 31 via a rubber pad 43 .
  • the setting of the unfinished turbine blade 1 ′ onto the jig 31 is completed. Since the spacing between the front edge of the front engagement face 27 f and the rear edge of the rear engagement face 29 f has been configured to be longer than the longitudinal length of the engagement member, the loaded state of the unfinished turbine blade 1 ′ on the jig 31 is further stabilized.
  • the manufacturing of the turbine blade 1 is finished by forming the engagement groove 25 s along the dovetail axial direction by the grinding. Since the front engagement face 27 f of the front engagement member 27 and the rear engagement face 29 f of the rear engagement member 29 are respectively configured to be substantially parallel to the dovetail axial direction, only the repulsive force, which is due to work resistance and is perpendicular to the dovetail axial direction, will occur on the front engagement face 27 f and the rear engagement face 29 f in a case where the engagement groove 25 s is formed along the dovetail axial direction by grinding. Therefore, substantially no repulsive force, which displaces the longitudinal direction of the dovetail, will occur.
  • the front engagement face 27 f of the front engagement member 27 and the rear engagement face 29 f of the rear engagement member 29 are respectively configured to be located slightly back from the virtual plane VF including one side of the platform 15 .
  • the end face of the front wall Wf and the end face of the rear wall Wr are respectively configured to be coplanar with the virtual plane VF. Therefore, the spacing between adjacent turbine blades 1 will not be widened locally when a number of turbine blades 1 are installed into the turbine disk 3 .
  • the front engagement member 27 and the front wall Wf are molded utilizing the thicker portion which is configured to be thicker in consideration to the strength of the front seal fin 21
  • the rear engagement member 29 and the rear wall Wr are molded utilizing the thicker portion which is configured to be thicker in consideration to the strength of the rear seal fin 23 . Therefore, the addition of the front engagement member 27 , the front wall Wf, the rear engagement member 29 and the rear wall Wr to the components of the turbine blade 1 does not cause any increase in a weight of the turbine blade 1 .
  • the present invention should not be limited to the description of the above embodiment of the invention, but it can be applicable in various modes through causing the appropriate conversion thereof, for example, an application of the turbine blade 1 to a turbine blade for a high-pressure turbine of the aircraft engine, etc.

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  • Engineering & Computer Science (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)
US10/743,091 2003-06-04 2003-12-23 Turbine blade Expired - Lifetime US7074012B2 (en)

Applications Claiming Priority (2)

Application Number Priority Date Filing Date Title
JP2003159175A JP4254352B2 (ja) 2003-06-04 2003-06-04 タービンブレード
JP2003-159175 2003-06-04

Publications (2)

Publication Number Publication Date
US20040247442A1 US20040247442A1 (en) 2004-12-09
US7074012B2 true US7074012B2 (en) 2006-07-11

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ID=33487461

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US10/743,091 Expired - Lifetime US7074012B2 (en) 2003-06-04 2003-12-23 Turbine blade

Country Status (6)

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US (1) US7074012B2 (ja)
EP (1) EP1631417B1 (ja)
JP (1) JP4254352B2 (ja)
CN (1) CN1798634B (ja)
DE (1) DE602004002697T2 (ja)
WO (1) WO2004108354A1 (ja)

Cited By (4)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070189695A1 (en) * 2006-02-10 2007-08-16 3M Innovative Properties Company Optical fiber loopback test system and method
US20130323071A1 (en) * 2012-06-01 2013-12-05 Pratt & Whitney Services Pte Ltd. Polishing assembly and method for polishing
US20140369844A1 (en) * 2012-02-02 2014-12-18 Snecma Optimisation of the bearing points of the stilts of vanes in a method for machining said vanes
US10989061B2 (en) 2016-10-21 2021-04-27 Siemens Energy Global GmbH & Co. KG Tip machining method and system

Families Citing this family (9)

* Cited by examiner, † Cited by third party
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US6863061B2 (en) 2003-01-15 2005-03-08 International Business Machines Corporation Row slicing method in tape head fabrication
US20060029500A1 (en) * 2004-08-04 2006-02-09 Anthony Cherolis Turbine blade flared buttress
EP1867436B1 (en) * 2005-03-09 2014-04-30 IHI Corporation Jig
US7536783B2 (en) * 2005-10-13 2009-05-26 Siemens Energy, Inc. Turbine vane airfoil reconfiguration method
US7503113B2 (en) * 2005-10-13 2009-03-17 Siemens Energy, Inc. Turbine vane airfoil reconfiguration system
RU2553049C2 (ru) 2011-07-01 2015-06-10 Альстом Текнолоджи Лтд Лопатка ротора турбины, ротор турбины и турбина
US10633985B2 (en) * 2012-06-25 2020-04-28 General Electric Company System having blade segment with curved mounting geometry
DE102013224199A1 (de) * 2013-11-27 2015-05-28 MTU Aero Engines AG Gasturbinen-Laufschaufel
CN113859764B (zh) * 2021-09-29 2023-04-07 中国航发动力股份有限公司 一种涡轮导向器组件导向叶片型腔的防护装置及制备方法

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US4638602A (en) * 1986-01-03 1987-01-27 Cavalieri Dominic A Turbine blade holding device
US4872812A (en) * 1987-08-05 1989-10-10 General Electric Company Turbine blade plateform sealing and vibration damping apparatus
JPH10196309A (ja) 1996-12-24 1998-07-28 United Technol Corp <Utc> タービンブレードプラットホームシール
US6017263A (en) 1996-04-30 2000-01-25 United Technologies Corporation Method for manufacturing precisely shaped parts
US6068541A (en) * 1997-12-22 2000-05-30 United Technologies Corporation Method for using a fixture enabling more accurate machining of a part
US6354803B1 (en) * 2000-06-30 2002-03-12 General Electric Company Blade damper and method for making same
US6786696B2 (en) * 2002-05-06 2004-09-07 General Electric Company Root notched turbine blade
US6842995B2 (en) * 2002-10-09 2005-01-18 General Electric Company Methods and apparatus for aligning components for inspection
US6855033B2 (en) * 2001-12-13 2005-02-15 General Electric Company Fixture for clamping a gas turbine component blank and its use in shaping the gas turbine component blank
US6857853B1 (en) * 2003-08-13 2005-02-22 General Electric Company Conical tip shroud fillet for a turbine bucket

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US5088894A (en) * 1990-05-02 1992-02-18 Westinghouse Electric Corp. Turbomachine blade fastening

Patent Citations (11)

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Publication number Priority date Publication date Assignee Title
US4638602A (en) * 1986-01-03 1987-01-27 Cavalieri Dominic A Turbine blade holding device
US4872812A (en) * 1987-08-05 1989-10-10 General Electric Company Turbine blade plateform sealing and vibration damping apparatus
US6017263A (en) 1996-04-30 2000-01-25 United Technologies Corporation Method for manufacturing precisely shaped parts
JPH10196309A (ja) 1996-12-24 1998-07-28 United Technol Corp <Utc> タービンブレードプラットホームシール
US5924699A (en) * 1996-12-24 1999-07-20 United Technologies Corporation Turbine blade platform seal
US6068541A (en) * 1997-12-22 2000-05-30 United Technologies Corporation Method for using a fixture enabling more accurate machining of a part
US6354803B1 (en) * 2000-06-30 2002-03-12 General Electric Company Blade damper and method for making same
US6855033B2 (en) * 2001-12-13 2005-02-15 General Electric Company Fixture for clamping a gas turbine component blank and its use in shaping the gas turbine component blank
US6786696B2 (en) * 2002-05-06 2004-09-07 General Electric Company Root notched turbine blade
US6842995B2 (en) * 2002-10-09 2005-01-18 General Electric Company Methods and apparatus for aligning components for inspection
US6857853B1 (en) * 2003-08-13 2005-02-22 General Electric Company Conical tip shroud fillet for a turbine bucket

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US20070189695A1 (en) * 2006-02-10 2007-08-16 3M Innovative Properties Company Optical fiber loopback test system and method
US20140369844A1 (en) * 2012-02-02 2014-12-18 Snecma Optimisation of the bearing points of the stilts of vanes in a method for machining said vanes
US20130323071A1 (en) * 2012-06-01 2013-12-05 Pratt & Whitney Services Pte Ltd. Polishing assembly and method for polishing
US9511469B2 (en) * 2012-06-01 2016-12-06 Pratt & Whitney Services Pte Ltd. Polishing assembly and method for polishing using a platform and barrier in a tumbling process
US10989061B2 (en) 2016-10-21 2021-04-27 Siemens Energy Global GmbH & Co. KG Tip machining method and system

Also Published As

Publication number Publication date
EP1631417A1 (en) 2006-03-08
JP4254352B2 (ja) 2009-04-15
DE602004002697T2 (de) 2007-10-04
WO2004108354A1 (en) 2004-12-16
CN1798634B (zh) 2010-06-09
JP2004360551A (ja) 2004-12-24
DE602004002697D1 (de) 2006-11-16
CN1798634A (zh) 2006-07-05
EP1631417B1 (en) 2006-10-04
US20040247442A1 (en) 2004-12-09

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